1. Field of the Invention
Embodiments of the present invention are related to injector assemblies, and more particularly, to injector assemblies having multiple manifolds for selective delivery of propellants.
2. Description of Related Art
Rocket engines provide thrust to a rocket, spacecraft, or other devices or vehicles by burning a mixture of a fuel, such as kerosene, methane, or hydrogen to list non-limiting examples, and an oxidizer, such as oxygen to list a non-limiting example, in a combustion chamber. The fuel and oxidizer are delivered to the combustion chamber by an injector assembly and are then atomized, vaporized, mixed, and combusted in the combustion chamber. The fuel and oxidizer are commonly referred to as propellants. The rocket engine may be throttled up to provide more thrust or throttled down to provide less thrust by increasing or decreasing the amount of propellants provided to the combustion chamber of the rocket engine. The individual propellants are often stored and initially delivered as liquids. Typically, injector assemblies of rocket engines and other applications are configured to operate with only one type of fuel, thus limiting the number of refueling possibilities that may further limit the use of the injector assembly.
As shown in FIG. 1, which illustrates a typical coaxial element injector assembly 10, the injector assembly comprises a manifold 12 for the oxidizer and a manifold 14 for the fuel, wherein each manifold may have a valve upstream of the manifold to control the amount of propellants provided to each manifold through the oxidizer inlet 12A and the fuel inlet 14A. The propellants flow from the manifold into the combustion chamber 16 through several individual injector elements 18 that comprise both an oxidizer flow passage and a fuel flow passage. The oxidizer flow passage and the fuel flow passage may define a coaxial arrangement to provide mixing of the oxidizer and fuel. An injector assembly may include only one injector element or up to several hundred injector elements in a large assembly.
In order to minimize the likelihood of poor performance or poor combustion stability, a typical injector element will have a pressure drop of 10% to 20% of the combustion chamber pressure during normal operation. Problems with injector assemblies often arise when the rocket engine is throttled up or down a relatively large amount which changes the pressure drop across the injector elements. This change in pressure drop is created by the change in flow through the injector elements, and the change in pressure is proportional to the square of the relative amount of propellant flow through the injector elements. For example, if the flow of the propellant is decreased to one half (½) of the original flow, the pressure drop is reduced by one fourth (¼) of the original pressure drop. Conversely, if the flow of the propellant is increased by three times, the pressure drop is increased by nine times.
If the pressure drop across the injector element is too low, atomization, vaporization, and mixing will be insufficient, thus leading to poor performance of the rocket engine. A low pressure drop across the injector element may also lead to combustion instability that may further lead to sudden failure of the rocket engine. If the pressure drop across the injector element is too high, an inordinate amount of energy is required to pump the propellant to the high pressure required to introduce flow into the injector assembly.
Therefore, a need exists for an injector assembly that maintains sufficient pressure drop when the injector is throttled a relatively large amount. In addition, the needed injector assembly would maintain good performance and protect from feed system coupled combustion instabilities without sacrificing the range of throttling available in conventional injector assemblies or exposing moving surfaces and dynamic seals to hot and/or corrosive reaction products.